49results about "Spacecraft vibration damping" patented technology

An apparatus comprises a core having a first surface configured for attachment to a surface of a structure, a face sheet located over a second surface of the core, a number of cavities within an interior of the core, and a number of ports for the number of cavities. The number of ports provides communication between the number of cavities within the interior of the core and the exterior of the core. The number of cavities and the number of ports are configured to reduce noise traveling through the core.

A method and apparatus for deploying a space structure. The space structure is secured to a base with release mechanisms that engage features of the space structure when the base is in a first position. The base is moved from the first position to a second position to deploy the space structure. The release mechanisms are moved at substantially a same time to disengage from the features and release the space structure from the base without imparting a transverse load to the space structure when the base moves from the first position towards the second position to deploy the space structure.

A method and system for calculating a control function for a structural system (10) that can be used to determine an appropriate control force to apply to an active member (18) within a stationary member (12) on the structural system (10). An active member (18) and a stationary member (12) are defined as a two-mass system in which the active member (18) and the stationary member (12) move in opposite directions. The stationary member (12) is mounted to an isolation subsystem (14) that is composed of six isolators (28) at multiple degrees of freedom. The isolation subsystem (14) is softer than the stationary member (12), active member (18) and a spacecraft surface (16) due to a damping element (32) of the isolation subsystem (16). The isolation subsystem (16) is mounted to the spacecraft (16) and decouples the spacecraft (16) from the stationary member (12) and thus the active member (18). An accurate control force for the active member (18) can be determined based upon the above structure (10).

Light-weight acoustic dampening launch vehicle fairing structures having high strength and stiffness and capable of filling to provide high mass density and rapidly unloaded at a predetermined time following launch, and methods of making them are provided. The panels include a monolithic body comprised of fibrous reinforcement material in a polymeric binder and comprise longitudinal chambers extending from the length of body. Fluid passage means are provided at both the top and bottom of the monolithic body and interconnected by the longitudinal chambers for filling and emptying said chambers with a fluid.

A magnetically damped mounting and isolation system with pointing capability. A payload is mounted to an isolator plate, a base plate is mounted to a satellite or other space vehicle, and the isolation system provides damping of all six degrees of freedom of isolator plate motion relative to the base plate. Three bidirectional magnetic dampers are connected between the isolator plate and the base plate and arranged to provide the required amount of temperature-independent damping. The bidirectional magnetic dampers can be connected to the base plate and the isolator plate in different configurations based on desired mass and natural frequency characteristics. Flexures which statically position the isolator plate are also designed to optimize normal modes of vibration. The isolation system may include a motion amplification feature to increase magnetic damping effectiveness, and the isolation system may also include active positioning of the payload relative to the satellite.

Launch lock assemblies with reduced preload are provided. The launch lock assembly comprises first and second mount pieces, a releasable clamp device, and a pair of retracting assemblies. Each retracting assembly comprises a pair of toothed members having interacting toothed surfaces. The releasable clamp device normally maintains the first and second mount pieces in clamped engagement. When the releasable clamp device is actuated, the first and second mount pieces are released from clamped engagement and one toothed member of each retracting assembly moves in an opposite direction relative to the other one toothed member of the other retracting assembly to define an axial gap on each side of the first mount piece.

A compartment for a space vehicle includes a pressurized structure having a structural wall, the structural wall having interior surfaces facing an interior of the compartment and exterior surfaces exposed to an external environment. An internal mounting structure for mounting a component is provided within the compartment, and mounting features support the internal mounting structure from the pressurized structure. The internal mounting structure is spaced away from the interior surfaces of the pressurized structure, and a thermal fluid is provided in the pressurized structure. The thermal fluid enables convective heat transfer between the component mounted on the internal mounting structure and the interior surfaces of the pressurized structure.

The invention discloses a device and method for satellite vibration isolation. The device for the satellite vibration isolation comprises an upper flange, a lower flange, a flexible hinge and at least two drive devices, wherein the upper flange is firmly connected to a satellite load; the lower flange is firmly connected to a satellite platform; the flexible hinge is disposed between the upper flange and the lower flange, and the flexible hinge is firmly connected to both the upper flange and the lower flange; and the drive devices are disposed between the upper flange and the lower flange, the drive devices are firmly connected to the upper flange and the lower flange, and the drive devices conduct radial rotation around the satellite load along with the lower flange according to drive control signals transmitted by a controller. The technical scheme provided by the invention has the advantages that a quiet and steady working environment can be provided for the satellite load; and a requirement of the satellite load for stability can be satisfied. In addition, the device for the satellite vibration isolation disclosed by the embodiment is characterized by a simple structure, high control accuracy, and low rigidness coupling, etc.

Embodiments of a low frequency isolation device are provided, as are embodiments of a spacecraft isolation system including a plurality of low frequency isolation devices. In one embodiment, the low frequency isolation device includes a three parameter isolator and a break frequency-reducing series spring mechanically coupled in series with the three parameter isolator and having a predetermined axial stiffness (K_{S AXIAL}) and a predetermined lateral stiffness (K_{S LATERAL}). The predetermined axial stiffness (K_{S AXIAL}) of the break frequency-reducing series spring is less than the predetermined lateral stiffness (K_{S LATERAL}) thereof.

The invention discloses a method for inhibiting vibration of satellite solar panel substrates by the aid of solid micro-thruster arrays. The method has the advantages that flexible satellite attitudekinetic equations with the solid micro-thruster arrays are created, joint active vibration controllers for the solid micro-thruster arrays and attitude flywheels are designed, accordingly, the vibration of the satellite solar panel substrates can be quickly inhibited, flexible satellites can be quickly stabilized after being disturbed, and the method is speedy.

The invention discloses a lightweight shock isolation device of an actuating mechanism on a satellite. The light shock isolation device of the actuating mechanism on the satellite comprises actuators, an upper platform and a lower platform. The upper platform and the lower platform are parallel to each other, and are fixedly connected with a vibration disturbance source and the satellite correspondingly. The three actuators are arranged between the upper platform and the lower platform, and the three actuators are hinged to the same point on the upper platform. The vertical projection of the mass center of the vibration disturbance source coincides with the hinged point of the three actuators on the upper platform. The three actuators are hinged to three points on the lower platform correspondingly, and connections of the three points form an equilateral triangle. The vertical projection of the mass center of the vibration disturbance source on the lower platform coincides with the center of the equilateral triangle. By mounting the actuators in different hinged ways, the mounting configuration of a control moment gyro is changed, and the singular difficulty of a control moment gyro group is overcome; through cooperation of a movable component, a shock isolation component and a sensor, the vibration disturbance problem of the actuating mechanism with high-speed rotors is solved; and the light shock isolation device of the actuating mechanism on the satellite has the characteristics of miniaturization and intellectualization.

Lightweight acoustic absorbers. The acoustic absorbers absorb energy during launch of a vehicle by a rocket, and are adapted to line the inside of a cavity of a fairing containing a vehicle. The absorbers comprise acoustic absorbing members, and a layer substantially surrounding the acoustic absorbing member which is substantially impermeable to contaminating particulates found within the fairing, and substantially impermeable to any contaminants which may be emitted by the acoustic absorbing member. A plurality of vents are cut in the layer, and a filter is placed over the vents to filter out any contaminants. The vents allow the gases within the fairing and the acoustic absorbing member to pass therethrough, thereby compensating for differential pressure.

Apparatus is provided for an acoustical composite construction especially designed for use in a launch vehicle fairing. The acoustical composite construction includes a low density core material having a first side and a second side. A first sheet of rigid face material is bonded to the first side with a visco elastic adhesive, and a second sheet of rigid face material is bonded to the second side with a visco elastic adhesive. The construction is then oven cured or air cured under vacuum bag pressure only.

An acoustic protection arrangement comprises acoustic protection modules, in particular acoustic absorber modules, wherein at least one of the acoustic protection modules is tuned to a different frequency, in particular has a different peak absorption frequency or acoustic frequency response, than at least another one of the acoustic protection modules. Furthermore, a kit of parts for assembling such an acoustic protection arrangement is provided, and a method for manufacturing such an acoustic protection arrangement is proposed. The acoustic protection arrangement can be implemented based on the kit of parts by selecting appropriate acoustic protection modules and combining them to form an acoustic protection arrangement that fulfils the specific acoustic and geometric requirements for an application, such as a launch vehicle type and/or launch mission. The acoustic protection arrangement can thus be adapted, in particular by tuning the proposed acoustic protection arrangement to achieve a desired overall level of acoustic attenuation.

The attenuation device of the present invention comprises a surface of revolution (11) of circular cross section, such that a set of slots (14) is distributed between the sides of the surface of revolution (11) and adapted to confine an elastic material (15) within the limits defined by the slots (14), generating a structural labyrinth for load passage from a lower interface of the surface of revolution (11) to an upper interface.

The invention discloses a variable-rigidity large-bearing vibration isolator suitable for vibration isolation of satellite equipment. The vibration isolator is mainly composed of a vibration isolationmodule, a rope module and a connecting module. The vibration isolation module is located in the vibration isolator, and the rope module is connected with the connecting module to form a bearing pathconnected with the vibration isolation module in parallel. The lower end and the lower end of the connecting module are connected with satellite equipment and a satellite structure correspondingly. The rope module has the advantages of being large in bearing capacity and high in rigidity under large deformation and can well adapt to the large-load environment in the launching process. The vibration isolation module has the advantages of being high in damping and low in rigidity under micro-deformation, and can well meet the efficient vibration isolation requirement in the on-orbit stage. The vibration isolator is simple in structure, good in space environment adaptability, high in reliability, light in weight and wide in application prospect.

The invention discloses a high-precision vibration isolation satellite and a control method thereof, and relates to the field of satellite vibration isolation control. The satellite comprises a satellite platform and a load platform, the satellite platform is connected with the load platform through an active vibration isolation control device, and a quality inspection and drag-free control deviceand an electrostatic suspension control device are arranged in the satellite platform; when the satellite is in an on-orbit flight process, the active vibration isolation control device is used for actively isolating high-frequency vibration between the satellite platform and the load platform; the drag-free control device is used for controlling the satellite platform to move along with the inspection mass, and the electrostatic suspension control device is used for controlling movement of the inspection mass. According to the high-precision vibration isolation satellite, full-band vibrationattenuation of the satellite is achieved, the load platform has the rapid high-precision attitude maneuver capacity, and the requirements of loads for the high microgravity level and high-precision rapid pointing are met.

The invention discloses an optical effective load passive vibration isolation system, and belongs to the technical field of spacecraft micro-vibration control. The problem that components of an existing vibration isolation platform are prone to be damaged under the condition that the ascending acceleration of a satellite is too high is solved. According to the system, the configuration is characterized in that three upper mounting supports are distributed and fixed to the side face of an upper platform, the positions of the three upper mounting supports are on the same circle, and the centralangle formed by every two upper mounting supports is 120 degrees; six eddy current damper mounting bases are fixed to a lower platform; six supporting legs are divided into three groups, the top endsof the two supporting legs in one group are fixed on the same upper mounting support, one eddy current damper is arranged at the bottom end of each supporting leg, and the eddy current dampers are fixed in eddy current damper mounting bases on the lower platform; and the plane where the two supporting legs in one group are located are perpendicular to the plane where the upper platform is locatedand the plane where the lower platform is located at the same time, the eddy current damper mounting bases at the bottoms of the three groups of supporting legs are distributed on the three edges of aregular triangle, and the eddy current damper mounting bases at the bottoms of one group of supporting legs are located on the same edge.

Provided is a non-adhesive type vibration reduction apparatus comprising: a support; a body; and an elastomer, wherein the elastomer is coupled in a non-adhesive manner in which an adhesive is not used, thereby preventing a separation of and reducing a damage of an elastomer which result from a weakening of an adhesive force.